CN113710875A

CN113710875A - Turbine engine blade, related turbine engine distributor and turbine engine

Info

Publication number: CN113710875A
Application number: CN202080028069.9A
Authority: CN
Inventors: 马修·西蒙; 让-吕克·巴查; 保罗·丹瑞; 莱安德烈·奥斯蒂诺
Original assignee: Safran Aircraft Engines SAS
Current assignee: Safran Aircraft Engines SAS
Priority date: 2019-04-03
Filing date: 2020-03-16
Publication date: 2021-11-26
Anticipated expiration: 2040-03-16
Also published as: US11795828B2; EP3947917B1; CN113710875B; FR3094743A1; FR3094743B1; WO2020201653A1; US20220170377A1; EP3947917A1

Abstract

A blade (13) for a turbine engine, arranged about an axis (X) and extending radially between a radially outer end (16) and a radially inner end (18), said blade (13) comprising at least one cooling cavity (24, 26) open to the radially outer end of the blade and to the radially inner end of the blade, said blade comprising at least one first tubular liner (34a, 36a) and at least one second tubular liner (34b, 36b), each liner being engaged in said cavity, the radially outer end of the first liner being open to the radially outer end of the blade, the radially inner end of the second liner being open to the radially inner end of the blade; the blade has a curvature such that the central axis of the blade has a radius of curvature of between 30 and 500mm, preferably between 30 and 100 mm.

Description

Turbine engine blade, related turbine engine distributor and turbine engine

Technical Field

The present invention relates to the field of aeronautical turbomachines, and more particularly to a blade for a turbine nozzle of a turbine engine, and a turbine engine comprising such a nozzle.

Background

As described in document FR2955145, the high-pressure or low-pressure nozzle of a turbine engine comprises in particular stator vanes held at both ends by an inner platform and an outer platform, these vanes defining a circulation flow path for the gases emitted from the combustion chamber. These vanes allow the gas exhausted from the combustion chamber to be directed to flow over the rotor blades of the turbine. These blades are hollow and comprise at least one or two cavities, in particular a leading edge cavity and a trailing edge cavity, one end of which opens to the outside of the flow channel.

These blades are exposed to high temperature combustion gases, which must be cooled to reduce thermal stresses. One solution is to use air from another component of the turbine engine, for example, the compressor. More specifically, upstream of the combustion chamber, relatively fresh air is taken at the outlet of the compressor stage. This air is injected into the cavity of the blade through both ends of the blade to internally cool the blade. The air then escapes into the flow passage through the holes made in the blade and communicates with the cavity of the blade and with the flow passage, the cooling air forming a protective film of fresher air flowing along the outer surface of the blade.

Furthermore, liners are typically embedded in the cavities of such blades. The pad includes a plurality of holes extending between the lands over the entire surface thereof. The gasket comprises a closed bottom wall at one end and opens to the outside of the flow channel on the same side as the cavity containing it. Relatively fresh air drawn upstream from the combustion chamber is injected into the liner through the platform. Fresh air enters the liner interior to cool the blades from the interior by impingement. However, integrating such a liner into an airfoil requires having a slightly curved blade airfoil in order to be able to insert the liner into the cavity. Such stresses limit the possibility of taking into account variations of the airfoil, in particular increasing the curvature (camber) of the blade.

Therefore, a device capable of solving the above technical problems is required.

Disclosure of Invention

The present disclosure relates to a blade for a turbine engine, the blade being configured to be disposed about an axis and to extend radially between a radially outer end and a radially inner end, the blade comprising at least one cooling cavity opening to the radially outer end of the blade and to the radially inner end of the blade, the blade comprising at least one first tubular liner and at least one second tubular liner, each liner being engaged in the cavity, the radially outer end of the first liner opening to the radially outer end of the blade, the radially inner end of the second liner opening to the radially inner end of the blade.

The ends of the blades and liners are, for example, the inner and outer radial ends when the blades are installed in a turbine. The blade is preferably hollow with a cavity extending on either side of the blade, each end. That is, the cavity of the vane corresponds to an aperture that extends through the entire height of the vane and opens at a radially inner end and a radially outer end. The end portions may for example be flanges provided around the blade at each end of the blade, to which flanges the ends of the liner are fixed.

According to the present disclosure, the first gasket is thus inserted into the cooling cavity from the radially outer end, and the second gasket is thus inserted into the cooling cavity from the radially inner end. In other words, when the gasket is mounted, the gasket is inserted into the cavity in two different mounting directions. This allows for two short pads rather than one long pad extending the entire length of the cavity. Thus, a significant curvature of the airfoil may be envisaged while maintaining efficient cooling of the blade over its entire height. This allows more freedom to improve the airfoil shape of, for example, the blades of the turbine nozzle, thereby increasing turbine efficiency.

In some embodiments, the radially outer end of the first liner is a first end secured to the radially outer end of the blade, the radially inner end of the second liner is a first end secured to the radially inner end of the blade, the first liner and the second liner each further including a second free end, the second free ends of the first liner and the second liner being opposite one another in the cavity for cooling the blade.

"opposite to each other" means that the second free ends are opposite to each other, or are in contact with each other, or are spaced apart from each other. In this second case, the second free ends are spaced apart from each other such that there is no separation therebetween. Thus, the second free ends of the pads do not overlap, making the mounting of each pad at one end of the blade easier.

In some embodiments, the first end of each liner includes an end flange that bears on the end of the blade.

In some embodiments, the flange is welded or brazed to the end of the blade.

This configuration allows to optimize the mounting of the gasket and its retention on the blade.

In some embodiments, the height of the first pad is different from the height of the second pad between the first and second ends of the pad along a direction radial to the axis, the height of each pad being less than 70%, preferably less than 60%, more preferably less than 50% of the length of the cavity in which the pad is engaged.

The height of the liners represents the length along their major axis, i.e., the radial direction when the blades are installed in a turbine engine. These height values of the pads with respect to the total height of the blade allow to obtain a greater curvature over at least a portion of the blade. For example, the blade has a greater curvature near its first end than near its second end.

In some embodiments, the height of the first pad is different than the height of the second pad.

This allows the height of each liner to be adjusted according to the desired airfoil shape of the blade.

In some embodiments, the radius of curvature of the central axis of the blade is between 30 and 500mm, preferably between 30 and 100 mm. The liners according to the prior art do not allow to design blades with a radius of curvature smaller than 90 mm. In particular, the turbine blades typically have a radius of curvature of between 90 and 500 mm. The liner according to the invention allows the manufacture of blades with a smaller radius of curvature.

In some embodiments, the central axis of the blade has a variable radius of curvature between the ends of the blade.

The central axis of the blade and liner represents the main axis along which the blade and liner extend. These radii of curvature are smaller than the radii of curvature present on the blades, which comprise a single liner extending over the entire height of the blade, and allow to improve the efficiency of the turbomachine on which the blades are mounted.

In some embodiments, the second end of each pad includes a bottom wall, and the distance between the bottom walls of each pad is less than 10 mm.

These values allow minimizing the portion of the cavity that does not include any liner, i.e., the space between the second ends of the first and second liners, thereby improving cooling of the blade.

In some embodiments, the second end of one of the first and second pads includes a protrusion protruding therefrom, and the second end of the other of the first and second pads includes an aperture into which the protrusion is inserted.

The protrusion may be inserted into the aperture, for example by pressure. This mode of connection allows the first and second pads to be secured together, thereby limiting the risk of the pads moving relative to each other.

In some embodiments, the blade comprises a leading edge cavity and a trailing edge cavity separated from the leading edge cavity by a wall, each of these cavities opening on a radially inner end and an outer end of the blade, a first liner engaged in each of the leading edge cavity and the trailing edge cavity, and a second liner engaged in each of the leading edge cavity and the trailing edge cavity.

That is, the blade includes four pads. Two first pads are inserted into the trailing edge cavity and the leading edge cavity from the radially outer ends, and two second pads are inserted into the trailing edge cavity and the leading edge cavity from the radially inner ends. The radii of curvature of the leading edge cavity and the trailing edge cavity may be different from each other.

In some embodiments, the radially inner and outer ends of the vanes are coaxial nozzle platforms configured to extend about an axis, the vanes extending between the platforms each having a flow passage face configured to define a gas circulation flow passage and a face opposite the flow passage face, the at least one cavity of the vane opening on the face opposite the flow passage faces of both platforms, an end of the first liner opening on the face opposite the flow passage face of the first platform, and an end of the second liner opening on the face opposite the flow passage face of the second platform.

The present disclosure also relates to a nozzle for a turbine engine, the nozzle being configured to be disposed about an axis and to extend radially between a radially inner platform and a radially outer platform, each platform comprising a flow passage face configured to define a fluid flow channel and a face opposite the flow passage face, the nozzle comprising at least one blade according to any one of the preceding embodiments, at least one cavity for cooling the blade opening onto an end face of the radially inner platform and an end face of the radially outer platform, a radially inner end of a first liner opening onto the end face of the radially inner platform, and a radially outer end of a second liner opening onto the end face of the radially outer platform.

The vanes of the present disclosure are stator vanes that extend radially between the platforms of the nozzle.

The present disclosure also relates to a turbine engine turbine comprising a nozzle according to the present disclosure.

The turbine may be a low pressure turbine or a high pressure turbine.

The present disclosure also relates to a turbine engine comprising a turbine according to the present disclosure.

Drawings

The invention and its advantages will be better understood by reading the following detailed description of various embodiments of the invention, given by way of non-limiting examples. The description makes reference to the accompanying drawings, in which:

FIG.1 illustrates an external perspective view of a turbine nozzle for a turbine engine according to the present disclosure;

FIG.2 illustrates an internal perspective view of the turbine nozzle of FIG. 1;

FIG.3 shows a perspective view of the liner of the present disclosure inserted into a hollow blade;

FIG.4 shows a perspective view of a liner inserted into a hollow blade according to the prior art;

fig.5 shows a perspective view of a modified example of the gasket of fig. 3.

Detailed Description

Fig.1 shows a segment 14 of a turbine nozzle for a high-pressure turbine engine, which nozzle can be segmented and comprises a blade ring or hollow stator blades 13 arranged between two coaxial platforms: an outer platform 16 and an inner platform 18. The

platforms

16, 18 may form an annular block, or include a plurality of ring segments arranged circumferentially end-to-end. They define a gas circulation flow channel 20 in which the vanes 13, which are evenly angularly distributed between the

platforms

16, 18, are located. Two blades 13 are shown. Each blade 13 includes a trailing edge cavity 26 opening to the exterior of the flow passage 20 through the platform 16 and the platform 18, and at least one leading edge cavity 24 opening to the exterior of the flow passage 20 through the platform 16 and the platform 18, the

cavities

24 and 26 being separated from one another by a wall 28. These cavities communicate with the flow passage 20 via a plurality of rows of

holes

30, 31 extending axially and/or radially along the blade 13 between the inner and

outer platforms

18, 16 to open into the flow passage 20. Thus, gas circulating from outside the flow passage 20 may enter the

cavities

24, 26, flow into the blade 13, and then be discharged through the

holes

30, 31 into the flow passage 20, thereby allowing cooling of the blade 13.

The first tubular liner 36a is inserted into the trailing edge cavity 26 and the first tubular liner 34a is inserted into the leading edge cavity 24 from the outer surface of the outer platform 16. "exterior surface" refers to the surface of the platform 16 opposite the flow channel 20.

In addition, a second tubular liner 36b is inserted into trailing edge cavity 26, and a second tubular liner 34b is inserted into leading edge cavity 24, from the outer surface of inner platform 18. "outer surface" refers to the surface of the platform opposite the flow channel 20.

Each

tubular liner

34a, 34b, 36a, 36b is hollow, may be made of a metal, such as a nickel-based or cobalt-based alloy, or of a composite material, and is perforated with a plurality of holes (not shown). The

first pads

34a, 36a each further include a flange 38a, the flange 38a being supported on the outer surface of the outer platform 16 and being secured thereto, such as by welding or brazing. The

second pads

34b, 36b each also include a flange 38b, with the flange 38b supported on the outer surface of the inner platform 18 and secured thereto, such as by welding or brazing.

The

first pads

34a, 36a of the leading edge cavity 24 and the trailing edge cavity 26 open to the exterior of the flow passage 20 through the platform (here the outer platform 16), while the

second pads

34b, 36b of the leading edge cavity 24 and the trailing edge cavity 26 open to the exterior of the flow passage 20 through the other platform (here the inner platform 18). Relatively fresh air from the compressor is directed on either side of the nozzle, i.e. outside of the outer platform 16 and outside of the inner platform 18. Thus, cooling air may enter the

liners

34a, 36a, 34b, 36b to cool the inner walls of the blades by impingement effect and then flow into the flow channels through the holes of the blades 13 to form a cooling film around each of them.

Fig.3 shows the arrangement when the

pads

34a, 34b, 36a, 36b are inserted into the

cavities

24, 26 of the blade 13, the latter being shown in broken lines, so that the pads are transparently visible. It should be noted in this figure that the

flanges

38a, 38b are not shown for simplicity of illustration.

The

cavities

24, 26 in the vane 13 have a depth H which corresponds approximately to the height of the vane 13 in the radial direction of the nozzle. Along the radial direction, the

first pads

34a, 36a have a height H1, and the

second pads

34b, 36b have a height H2. According to this embodiment, heights H1 and H2 are substantially equal. However, this example is not limiting, and the heights H1 and H2 may be different as long as H1 and H2 remain less than 70%, preferably less than 60%, more preferably less than 50% of the H value.

The first trailing edge pad 36a includes a bottom wall 361a that closes off the pad 36a at an end opposite the end that opens to the exterior of the outer platform 16. Likewise, the second trailing edge pad 36b includes a bottom wall 361b, the bottom wall 361b closing the pad 36b at an end opposite the end that opens to the exterior of the inner platform 18. The

bottom walls

361a and 361b face each other within the cavity 26, either contacting each other or alternately spaced apart from each other by a distance D of less than 10 mm.

Similarly, the first leading edge pad 34a includes a bottom wall 341a that closes the pad 34a at an end opposite the end that is open to the exterior of the outer platform 16. Likewise, the second leading edge pad 34b includes a bottom wall 341b, the bottom wall 361b closing the pad 34b at an end opposite the end that opens to the exterior of the inner platform 18. The

bottom walls

341a and 341b face each other within the cavity 24, contacting each other or alternately spaced apart from each other by a distance D.

According to this embodiment, the

first cushions

34a and 36a are inserted from the outside of the outer platform 16, while the

second cushions

34b and 36b are inserted from the outside of the inner platform 18. Thus, the curvature of the vane 13 may be greater than a configuration using a single pad for each

cavity

24, 26 over the entire height H. Fig.4 shows this situation according to the prior art, where the blade 13 has a curvature similar to that of the blade 13 in fig.3, and illustrates the difficulty of inserting a gasket in this situation.

Fig.5 represents a modified example of the embodiment of fig.3, in which the

pads

36a and 34b each comprise a stud 40, which may be hemispherical with a radius of less than 10mm, and which protrudes radially from the

bottom walls

361a and 341b, respectively. These studs 40 are configured to ensure a gap between the

bottom walls

341a and 341 b. The radius of the stud 40 may preferably be less than 0.1mm or less than the distance D between the

bottom walls

341a and 341b to limit mechanical stress between the two gaskets. Alternatively, the stud 40 may be inserted into an aperture 42 formed in the opposing bottom wall (i.e.,

walls

361b and 341a, respectively). This arrangement allows the pads to be positioned relative to each other and restricts movement relative to each other. It should be noted that the location of the stud 40 and the aperture 42 is not limiting, and the stud 40 can be disposed on, for example, the

walls

361b and 341a or 361a and 341a, the aperture can be disposed on the

walls

361a and 341b or 361b and 341b, or other possible combinations.

Although the present invention has been described with reference to specific exemplary embodiments, it will be apparent that modifications and variations can be made to these examples without departing from the general scope of the invention as defined in the claims. In particular, individual features of different illustrated/referenced embodiments may be combined in additional embodiments. The specification and drawings are, accordingly, to be regarded in an illustrative rather than a restrictive sense.

Claims

1. A blade (13) for a turbine engine, configured to be arranged around an axis (X) and to extend radially between a radially outer end (16) and a radially inner end (18), the blade (13) comprising at least one cooling cavity (24, 26) opening to the radially outer end of the blade and to the radially inner end of the blade, the blade (13) comprising at least one first tubular liner (34a, 36a) and at least one second tubular liner (34b, 36b), each liner being engaged in the cavity (24, 26), the radially outer end of the first liner opening to the radially outer end (16) of the blade and the radially inner end of the second liner opening to the radially inner end (18) of the blade, the blade (13) having a curvature such that the central axis of the blade (13) has a diameter of between 30 and 500mm, preferably between 30 and 100 mm.

2. The blade (13) of claim 1, wherein the radially outer end of the first liner (34a, 36a) is a first end fixed to the radially outer end (16) of the blade, the radially inner end of the second liner (34b, 36b) is a first end fixed to the radially inner end (18) of the blade, the first liner (34a, 36a) and the second liner (34b, 36b) each further comprising a second free end, the second free ends of the first and second liners being opposite to each other in the cavity (24, 26) for cooling the blade (13).

3. A blade (13) according to claim 2, wherein the first end of each gasket comprises an end flange (38a, 38b) bearing on an end (16, 18) of the blade, the flange (38a, 38b) being welded or brazed to the end.

4. Blade (13) according to claim 2 or 3, wherein the height (H1) of the first gasket (34a, 36a) differs from the height (H2) of the second gasket (34b, 36b) between its first and second ends, along a direction radial to the axis (X), the height of each gasket being less than 70%, preferably less than 60%, more preferably less than 50%, of the length of the cavity (24, 26) in which it is engaged.

5. A vane (13) as claimed in any one of claims 2 to 4 wherein the second end of each liner comprises a bottom wall (361a, 361b), the distance (D) between the bottom walls of each liner being less than 10 mm.

6. The blade (13) according to any of claims 2 to 5, wherein a second end of one of the first and second pads (34a, 36a, 34b, 36b) comprises a protrusion (40) protruding therefrom, a second end of the other of the first and second pads comprises an aperture (42), the protrusion (40) being inserted into the aperture (42).

7. A blade (13) according to any of claims 1-6, wherein the blade (13) comprises a leading edge cavity (24) and a trailing edge cavity (26) separated from the leading edge cavity by a wall (28), each of these cavities (24, 26) opening on a radially inner and outer end (16, 18) of the blade, a first liner (34a, 36a) being engaged in each of the leading and trailing edge cavities (24, 26), and a second liner (3ba, 36b) being engaged in each of the leading and trailing edge cavities (24, 26).

8. A nozzle (14) for a turbine engine, the nozzle being configured to be arranged around an axis (X) and to extend radially between a radially inner platform (16) and a radially outer platform (18), each platform (16, 18) comprising a flow passage face configured to define a fluid flow passage and a face opposite the flow passage face, the nozzle comprising at least one blade (13) according to any one of the preceding claims 1 to 7, at least one cavity (24, 26) for cooling the blade (16) opening on an end face of the radially inner platform (16) and an end face of the radially outer platform (18), a radially inner end of the first liner opening on an end face of the radially inner platform (16), a radially outer end of the second liner opening on an end face of the radially outer platform (18).

9. A turbine engine comprising a nozzle (14) according to claim 8.